This invention relates to thermal protective materials (TPMs) for the aerospace industry and more particularly, to a reinforced carbon composite material which has variable substrate density prior to impregnation, which is impregnated with a silicon based ablative resin which is cured and manufactured to form structural configurations which are useful for mounting on the exterior surface of a structure to be protected by the TPM and the method of making same.
During reentry into the atmosphere, a vehicle is subjected to extreme thermal conditions. As the vehicle contacts the atmosphere at very high speeds, frictional forces release high levels of thermal energy which can raise the temperature to levels which are destructive to the outer shell. To protect the vehicle from high temperatures and wind shear, the vehicle""s outer shell is typically covered with TPMs, which act as insulators and are designed to withstand these extreme thermal conditions.
Carbon-carbon (Cxe2x80x94C) composites are one class of TPM which have been employed under such conditions with proven effectiveness. The success of a particular TPM requires that the system have sufficient mechanical strength at high temperatures, produce endothermic reactions upon decomposition, and have a high surface emissivity.
In its simplest form, a carbon-carbon composite is manufactured by combining carbon fibers with an organic resin, usually a high carbon yield epoxy or phenolic resin, and the resulting carbon fiber and resin matrix cured to achieve a three dimensional structure such as a tile, billet or other object. The matrix has a density, a void volume and a degree of mechanical strength.
The carbon fiber and resin matrix is then subjected to a high temperature treatment which decomposes the resin matrix to pure carbon, a process called charring or carbonization. Charring changes the resin coating from an organic resin to free carbon which coats the carbon fibers and partially fills the void spaces of the matrix with free carbon. The TPM may be subjected to several charring cycles, a process known as densification. The result of densification is to create a more rigid substrate, with a decreased void volume. The char surface of the substrate has a high temperature structural capability, which is a desirable characteristic.
Conventional Cxe2x80x94C composites are manufactured in such a way so as to produce a highly filled and rigidized structure with a minimum of porosity. There are many ways for Cxe2x80x94C materials to be densified including infiltration with petroleum pitch, impregnation with phenolic or other organic resins, or carbon vapor infiltration (CVI) using low molecular weight hydrocarbons such as methane. Any substance used for densification should have a high carbon char yield. Repeated cycles of impregnation and carbonization are required to first infuse the material with the carbon materials and then to heat them to a sufficiently high temperature (generally above 500xc2x0 C.) to char the infiltrant and to create porosity for further densification cycles. A typical density range for a Cxe2x80x94C composite with 5% porosity is approximately 1.6 to 1.8 g/cc, depending on the infiltrants and carbon fibers used in the composite.
The use of Cxe2x80x94C composite TPMs on long duration, high altitude hypersonic reentry vehicles exhibit, however, some characteristics which can severely restrict mission performance. A major limitation of these materials is that they are subject to oxidation at extreme thermal conditions. The oxidation that these TPMs experience during long duration reentry can result in large shape changes to the vehicle aero-shell. Shape changes that adversely affect the mechanical strength and aerodynamics of the vehicle are unacceptable. To compensate for the loss of mechanical or structural integrity, which can lead to shape changes, typically the thickness of the material is increased. Increasing the thickness, however, adds unacceptably to the weight and volume of the vehicle, thus reducing the payload capacity and increasing cost.
While the Cxe2x80x94C class of TPMs make them good candidates for aerospace applications due to their excellent high temperature structural properties oxidation shape changes can still be a problem. To address this, extensive efforts have been expended on oxidation resistant coatings for Cxe2x80x94C composites with, however, limited success. The coatings developed to date are restricted to temperature levels generally below those experienced during reentry into the atmosphere, or in other high temperature applications. Also, coating costs and durability (durability in the form of handling microcracking, the occurrence of pinholes, particle impacts and damage from ground handling) are serious issues when one is considering coatings for use on Cxe2x80x94C composite TPMs.
Ablation technology employs several mechanisms to manage the high levels of thermal energy released during reentry. Three of these are the vaporization and decomposition (pyrolysis) of the resin and subsequent transpirational cooling at the boundary layer. All of these processes absorb heat. Producing large amounts of gas is one measure of an ablation based system""s ability to absorb heat. The production of gas can also be increased by impregnating the Cxe2x80x94C substrate with an organic material specifically designed to vaporize and pyrolyze upon exposure of the system to high heat loads. Materials used in these passive transpiration systems, known as coolants, include materials such as polyethylene or epoxy, acrylic or phenolic resins.
Under such a system, there is created within the material a pyrolysis zone, where the resin and any supplemental coolants present are heated to temperatures where the organic materials decompose. The effect is the absorption of heat and the creation of additional carbon which can remain in the pyrolysis zone and/or be deposited on the carbon fibers and within the void volume of the substrate. Thus, the Cxe2x80x94C ablator""s final weight and ability to absorb heat are directly related to the amount of available resin in the Cxe2x80x94C composite prior to reentry.
At the surface of the Cxe2x80x94C ablator, heat is reradiated due to the refractive properties of the carbon substrate. In addition, the gasses produced in the pyrolysis zone within the Cxe2x80x94C ablator are released to the surface at a relatively cool temperature when compared to the conditions at the surface. This effect, known as pyrolysis gas transpiration, provides cooling at the surface of the TPM. The disadvantages of the passive transpiration systems described herein include the high overall density of the material and the high internal pressure cause by the sudden buildup of gasses within the material. Ablation systems which can create and then release large volumes of gas thus exhibit a greater capacity to absorb and dissipate the heat of reentry.
In this regard, the structure of the Cxe2x80x94C substrate is important to the overall effectiveness of the ablator. The void volume can be filled with a resin or other coolant to provide the raw material for production of gasses. In addition, methods of construction of the substrate can allow for greater transpiration pathways for release of the gasses. Systems which generate large volumes of gas over a short period of time also generate high internal pressures. Such pressure causes internal cracking in the substrate (microcracks) and also spalling at the surface. These effects are destructive to the mechanical integrity of the system and can lead to system failure. Therefore, improved transpiration pathways also protect the system from the effects of this internal pressure.
U.S. Pat. No. 5,635,300 to Kostikov, et. al., describes an advancement in the art of Cxe2x80x94C or ceramic ablators through the introduction of silicon based resins to the Cxe2x80x94C substrate. Upon decomposition and subsequent exposure to the very high temperatures at the surface, the silicon resin reacts with the carbon substrate to form a silicon carbide (SiC) coating on those fibers experiencing the high temperature conditions. The formation of SiC is more resistasnt to oxidation than carbon, and thus acts to strengthen the carbon substrate by forming a SiC skeleton in the areas of extreme temperature. When prolonged conditions of high temperature and wind shear at the surface lead to loss of SiC, the newly exposed carbon substrate undergoes further reaction to form new SiC, thus regenerating the protective skeleton.
The SiC layer which forms over the fibers of the carbon substrate in the interior of the ablator has a different coefficient of thermal expansion (CTE) than the carbon itself. The result is that when the system is subjected to temperature changes, the SiC coating within the carbon substrate forms microcracks. These cracks form passageways for the entry of air, which leads to oxidation of the carbon substrate, with resultant loss of strength and integrity of the ablator.
In Kostikov, a carbon-SiC substrate is created by obtaining a carbon-plastic preform composed of carbon fibers and a thermosetting resin binder and heat treating to form a coke matrix reinforced by carbon fibers. The coke matrix is densified by infiltrating with pyrolytic carbon and heat treating the preform at 1900 to 2000 deg. C. According to this invention, pore channels are formed following crystalization of the carbon deposited upon the matrix. Densification is followed by treatment with silicon which forms a SiC skeleton in the pore spaces of the composite. The carbon fibers can be in the form of a woven fabric or woven substrate.
U.S. Pat. No. 5,672,389 to Tran, et. al., discloses a low density ceramic ablator which employs a fibrous ceramic substrate which has, prior to impregnation with a resin matrix, a density of about 0.15 to 0.2 g/cc. Tran includes carbon fibers within a definition of the term ceramic. The ceramic substrate is impregnated with a low viscosity solution containing of an organic resin in a solvent. The excess infiltrant is removed, followed by removal of the solvent under vacuum, leaving resin-coated fibers, and a substrate having an average density of from 0.15 to 0.4 g/cc. Tran discloses that the resulting ablator may have either a uniform distribution of resin on the ceramic fibers, or a non-uniform distribution. The non-uniform distribution has the benefit of achieving the necessary degree of ablation at the outer surface, while being lightweight at the inner surface, where extreme temperatures are not experienced.
In addition, at the oxidizing conditions at the surface, silicon reacts with atmospheric oxygen to form a coating of silicon oxide (SiO2), which manifests itself as a glassy layer on the outer surface of the ablator. This SiO2 plus free carbon and free SiC mixture has a high surface emissivity which improves the ability of the material to radiate heat from the surface due to convection and re-radiation from the carbon substrate.
U.S. Pat. No. 5,965,266 to Goujard, et. al. discloses a carbon-SiC (Cxe2x80x94SiC) composite TPM which has a self-healing mechanism for in-situ repair of the Cxe2x80x94SiC matrix. The matrix is heat treated to form SiC and boron carbide (BC) over the Cxe2x80x94SiC matrix. The SiC layer improves the mechanical strength of the system. However, due to the difference in CTE, the matrix experiences destructive cracking when exposed to the temperature changes experienced during reentry. These cracks create pathways that allow the entry of air, which causes oxidation of the Cxe2x80x94SiC matrix, thus mechanically weakening the structure of the TPM.
The invention of Goujard provides available free silicon and boron as glass precursors which react with the exposed carbon under the high temperatures and oxidizing conditions of reentry. The glass precursors to form a self healing glass layer within the cracks, closing the pathway for internal oxidation of the substrate.
In addition to the strength and density of an ablation system, the manner in which the material is mounted onto the aero-shell, and the incorporation of additional layers of insulation also have bearing of the success of the TPM. U.S. Pat. No. 3,152,548 to Schwartz discloses a system whereby a series of coiled wires are attached to the aero-shell and the ceramic TPM is mounted onto the metal coils, thus creating a space between the aero-shell and the ceramic TPM. This space is filled with a pliable thermal insulating material, thus providing added insulation protection to the aero-shell. The patent discloses that the use of coiled wire compensates for differences in thermal expansion between the metal aero-shell and the ceramic insulator.
It is therefore a principal object of the invention to provide a thermal protection material (TPM) which is a carbon-carbon (Cxe2x80x94C) ablator, and which is of relatively low cost, low density, high mechanical strength, and which offers a high degree of protection from oxidation. The present invention allows the TPM to be manufactured in a manner in which the variables of strength, weight and heat absorption can be varied across the thickness or the length of the TPM so as to achieve an optimal balance of these variables at the lowest attainable cost. The Cxe2x80x94C ablator of the present invention also provides a structure with passageways, which allow for improved rates of transpiration of gasses produced. The Cxe2x80x94C ablative TPM of this invention also includes methods of construction which allow for new and useful configurations of the Cxe2x80x94C ablative TPM so that insulation material may be incorporated between the TPM and the vehicle aero-shell.
The Cxe2x80x94C substrate of the present invention is a three dimensional object which can be woven or non-woven. The density of the fibers increases across the thickness of the TPM, thus increasing the strength of the substrate in that direction. Fiber density can be varied by varying either the weaving method or the type of fabric used (i.e. woven, non-woven, knitted or braided fabrics). In addition, the invention may include needling of the fabric, which increases the interconnection of the fibers across the structure""s thickness (the z direction). Needling also serves to increase the porosity in the z direction, thus providing improved pathways for the transpiration of ablation gasses produced under high temperature conditions, such as during reentry. Improved weaving methods which can create three-dimensional structures may also be employed to vary fiber density in the z direction, and to increase transpiration rates of the ablator.
The Cxe2x80x94C ablator of this invention is impregnated with an organic resin having a high carbon yield, and the resulting matrix is cured. The resulting coated substrate is subjected to one or more charring cycles to densify the substrate. After the charring cycles, the system is treated with a silicon based ablative resin and cured.
Inherent in this invention is the use of a silicon based ablative resin. Silicon based resin is available below the surface of the composite, and when heated, flows to the surface to react with carbon to produce SiC. Thus, while the ablative resin serves as a coolant to the system, the chemical reactions which occur at high temperatures within the ablator also serve to provide mechanical strength to the Cxe2x80x94C substrate by creating an oxidation resistant SiC coating.
During reentry, the high temperatures are sufficient to oxidize the carbon substrate. This leads to recession of the surface of the ablator, resulting in a loss of mechanical strength and, consequently, changes to the shape of the vehicle surface. These shape changes can negatively affect the aerodynamics of the vehicle, which is unacceptable. The high temperature of the system during reentry creates SiC from a reaction of the silicon with the carbon of the substrate. During reentry, then, part of the char layer is oxidized. As this occurs, carbon is replaced by SiC, which provides a protective coating which resists oxidation. Furthermore, as recession proceeds at the surface of the ablator, the exposed carbon substrate further reacts with the silicon to form a layer of silicon carbide at the affected area.
In addition, at high temperatures, the silicon in the gasses produced by the pyrolysis of the resin react with the oxygen in the atmosphere at the ablator""s surface to produce silicon dioxide (SiO2), along with a mixture of free carbon and SiC. This mixture can be highly transmissive. With further oxidation, the concentration of SiO2 increases at the surface, providing oxidation protection to the subsurface carbon and SiC.
Another aspect of this invention is that the formation of SiC does not occur until the system is subjected to the high temperatures of reentry. This feature of creating the SiC matrix in-situ avoids the destructive effects of microcracking which occurs when a Cxe2x80x94C substrate and SiC matrix is subjected to high temperature changes and/or mechanical stresses.
What has been developed is a unique approach to forming a low cost Cxe2x80x94C composite with an in-situ method for providing durable oxidation protection. Not only is this a lower cost approach, but it offers improved ablation resistance compared with the conventional Cxe2x80x94C composites used to date. A lower density Cxe2x80x94C (1.3 to 1.5 g/cc) is used. This results in significant cost savings due to fewer numbers of densification cycles required. This material is infiltrated with a silicon based ablative resin (such as an RTV, manufactured by General Electric Corp.) using a resin transfer molding process (RTM).
The RTM process involves evacuating the sample of air in a closed mold and pressure impregnating the sample with RTV to fill in the available pores of the substrate. Since the invention requires that RTV materials are stored within the substrate, to be available for protection during reentry heating, provisions for this storage area are provided in the form of increased void volume. Furthermore, the Cxe2x80x94C substrate must be designed and fabricated so that this storage volume and distribution is predictable. This is important to control the amount of RTV since too much RTV can create high internal gas pressure during high temperature exposure. Not enough RTV results in a loss of the protective effects of the ablative system during the reentry phase of flight.
Other embodiments of the invention include a Cxe2x80x94C substrate which is created by a three dimensional weaving process known as multilayer interlock braiding. This weaving method can be used to produce a three dimensional carbon fiber substrate with the desired gradation in fiber density in a direction. The three dimensional woven object has great strength in the z direction and avoids the problems experienced with two dimensional fabrics which can have less integrity and interconnection across the layers of the composition.
Multilayer interlock braiding is a technique which allows for an interconnected three dimensional braid which can be formed into a three-dimensional form. The technique can attain a 3-D structure with variation in fiber density in the z direction. The resulting 3-D substrate has increased strength in the z-direction and allows for improved pathways for gas transpiration in the z-direction. The resulting substrate can be needled for additional interconnection and transpiration pathways. Multilayer interlock braiding is described in an article entitled 3-D Braided Composites, Design and Applications, Brookstein, D. (Albany International Research Co., Sixth European Conference on Composite Materials, September, 1993), the disclosure of which is incorporated herein by reference.
An additional advantage of the current invention, not achievable by any compositions in the art, is the ability to control not only the density versus the strength of the substrate, but also the amount of ablative resin available to be pyrolyzed and therefore available as coolant. The current invention has variable density and void volume across the thickness of the TPM. The areas with higher void volume contain greater amounts of RTV available for ablation during reentry. Therefore, the cooling capacity of the TPM varies depending upon the density of the Cxe2x80x94C substrate. The external layers may contain a higher volume of coolant, while the interior layers can exhibit a higher mechanical strength.
To provide material systems that meet the above requirements for porosity and outgassing, several candidate manufacturing techniques are available. These include an open woven knit structure, since the loops inherent with the knitting process provide natural pockets of porosity available for RTV storage. Another concept is a woven structure with intentional spaces between adjacent yarns for creating the volume needed for RTV storage. Another concept is a multi-layer woven structure using a Jacquard type weaving machine. The weave architecture in such a structure can easily be tailored to provide volume for RTV storage. Another concept that offers the lowest cost option is a non-woven preform. Such a preform can be manufactured with preformed orientation in a layered construction. In addition, this concept as well as the other concepts mentioned can benefit from a needling process for added structural integrity.
There are a number of ways to incorporate outgassing paths for the structure. One way is to needle the preform prior to the Cxe2x80x94C densification process. This process pierces the face of the woven preform in a regular or a tailored pattern or grid. This needling process causes a percentage of the fibers being pierced to align along the needling direction, creating a through thickness fiber component. This not only produces paths by which the gasses migrate out of the component, but the added through-thickness reinforcement increases interlaminar mechanical properties.
Another mechanism for providing through thickness gas paths is T-forming. T-forming is a method by which fibers are inserted directly into the preform. T-forming is disclosed in U.S. Pat. No. 6,103,337 assigned to Albany International Corp., Techniweave Division, entitled, Fiber Reinforced Structures and Method of Making Same, the disclosure of which is incorporated herein by reference. With this method, the T-forming spacing depth of penetration, and the orientation can be controlled. T-forming may also be a method selected for mechanically attaching the outer protection layer to the support elements to produce three-dimensional structural components.
By employing T-forming, the material concepts described can be tailored to the particular application requirements. The TPM can be manufactured in configurations which are structurally capable of withstanding the thermally induced structural loads and the aerodynamic loads of reentry and maneuvering. The material system can be designed to effectively transfer the loads while not act as a heat path from the aero-shell.
Another improvement of this invention involves the improved methods of attaching insulation material between the ablator and the vehicle""s outer shell. The ablator, being made from a carbon fiber substrate, can be formed into advantageous configurations. These structural features can take the form of T-formed ribs and stiffeners, Cxe2x80x94C honeycombs, integrally woven ribs, corrugated Cxe2x80x94C and other advantageous forms. The space produced by installing the ablator in the form of a corrugated, T-joined or similar configuration is filled with an insulator material to add additional heat protection to the system.
Another embodiment of this invention is to fabricate a carbon fabric tape material with the silicon bearing RTV impregnated into the fabric surface. This material can then be laminated using heat and pressure to form a structural, fiber reinforced component with the silicon protection scheme already in place. This is a process that would require no Cxe2x80x94C processing.